1. Field of the Invention
This invention relates to fuel systems for use in gas turbine engines.
2. Summary of the Prior Art
In order to operate efficiently, gas turbine engines must be run at very high temperatures. As one would expect, these temperatures get particularly high in a combustor section of the engine where engine fuel is burned. High combustor temperatures are necessary in order to fully ignite the fuel and, additionally, to derive the maximum amount of energy available from the burning fuel. As the fuel is ignited, it combines with high pressure air to form high-temperature, high-pressure combustion gases. These gases are utilized downstream of the combustor by a turbine section where the kinetic energy of the gases is transformed into useful mechanical energy. Under basic thermodynamic principals, increasing the temperature and pressure of the combustion gases increases the amount of mechanical energy produced.
Because of the necessarily high combustor temperatures, an engine fuel system must be provided that is capable of safely and reliably supplying a continuous flow of fuel to the combustor during high temperature engine operation. Typically, in the present state of engine development, fuel systems are subjected to temperatures in excess of 800.degree. F. (426.67.degree. C.). The Federal Aviation Authority (FAA) requires that commercial engine fuel systems undergo a flame endurance test to show that a particular engine's fuel system is capable of safe operation in this harsh environment. At least one aircraft engine manufacturer, namely the assignee of the subject invention, additionally requires that the fuel system must be capable of carrying any fuel leakage overboard in the event a fitting or a line in a primary flowpath should fail.